Low hub-to-tip ratio fan for a turbofan gas turbine engine

ABSTRACT

A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor. Each fan blade defines a leading edge. A hub radius (R HUB ) is the radius of the leading edge at the hub relative to a centerline of the fan. A tip radius (R TIP ) is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan. The ratio of the hub radius to the tip radius (R HUB /R TIP ) is at least less than 0.29. In a particular embodiment, this ratio is between 0.25 and 0.29. In another particular embodiment, this ratio is less than or equal to 0.25.

TECHNICAL FIELD

The present invention relates to turbofan engines and more particularly to fans for such engines having low hub to tip ratios.

BACKGROUND

Most gas turbine engine fans are composed of a central hub onto which a plurality of separately formed fan blades are secured. Integrated bladed rotor (IBR) fans are known for their relative lightness and therefore are desirable, however known IBR fans cannot be formed having a low hub to tip radius ratio because of limitations in manufacturing capabilities. Such a low hub to tip radius ratio is however desirable because it means the maximum diameter of the fan can be reduced without negatively effecting performance. Reducing the overall diameter of the fan reduces weight and improves the efficiency of the fan.

Therefore, while the advantages of reducing the ratio of the radius of the hub to the radius of the tip are well appreciated in terms of reducing the specific flow of air entering the leading edge of the fan, attempts to date to reduce the specific flow by reducing this ratio have not been readily possible, particularly for IBR fans. Attempts to manufacture an integrated bladed rotor (IBR) fan with a low hub to tip ratio have not been successful because of the lack of space for machine tools between the roots of the blades when the hub is also reduced in size.

SUMMARY

There is therefore provided a fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of radially extending fan blades integral with the hub to form an integrally bladed rotor, each fan blade having a leading edge, a hub radius (R_(HUB)) which is the radius of the leading edge at the hub relative to a centerline of the fan, and a tip radius (R_(TIP)) which is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan, and wherein the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is at least less than 0.29.

In a particular embodiment, the ratio R_(HUB)/R_(TIP) is less than or equal to 0.25.

In another particular embodiment, the ratio R_(HUB)/R_(TIP) is between 0.25 and 0.29.

There is also provided a method of manufacturing an integrally bladed rotor fan for a turbofan gas turbine engine, comprising: forming a rotor hub preform defining a hub radius and having at least a number of root stubs radially spaced apart on a periphery of the rotor hub perform; providing blade airfoils having a length such that a ratio of the hub radius to a tip radius of the blade airfoils, once mounted to the hub, is at least less than 0.29; and subsequently fastening the blade airfoils to the root stubs to form fan blades integrally formed with the hub resulting in an integrally bladed rotor fan having a hub to tip radius ratio of at least less than 0.29.

There is further provided a turbofan gas turbine engine comprising a fan upstream of at least one compressor, the fan having a rotor hub and a plurality of substantially radially extending fan blades integral with the rotor hub to form an integrated bladed rotor, each said fan blade having an airfoil defining a leading edge and defining a tip radius (R_(TIP)) which is the radius of a tip of the fan blade at the leading edge, the rotor hub defining a hub radius (R_(HUB)) which is the radius of the hub at the blade leading edge, and wherein a ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is at least less than 0.29.

BRIEF DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures in which:

FIG. 1 is a schematic cross-sectional view of a gas turbine engine having a fan in accordance with the present disclosure; and

FIG. 2 is a partial axial cross-sectional view of an embodiment of the fan of the present disclosure.

DETAILED DESCRIPTION

FIG. 1 illustrates a turbofan gas turbine engine 10 generally comprising in serial flow communication, a fan assembly 12 through which ambient air is propelled, and a core 13 including a compressor section 14 for pressurizing the air, a combustor 16 in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, and a turbine section 18 for extracting energy from the combustion gases. A centerline main engine axis 13 extends longitudinally through the turbofan engine 10.

The fan 12 propels air through both the engine core 13 and the bypass duct 22, and may be mounted to the low pressure main engine shaft 11. The fan 12 includes a plurality of radially extending fan blades 20 and a central hub as will be seen, which hub has a nose cone 22 mounted thereto to protect the hub. As will be described in greater detail below, the fan 12 is an integrally bladed rotor (IBR), wherein the fan blades 20 are integrally formed with the central hub that is fastened to the low pressure (LP) engine shaft 11 for rotation therewith.

Referring now to FIG. 2, the IBR fan 12 comprises a plurality of fan blades 20 integrally formed with, and substantially radially extending from, a central fan hub 36 which is mounted to an engine shaft, such as the low pressure shaft 11, by means of one or more hub support portions 38 which are also integrally formed with the hub. Each of the blades 20 defines an airfoil 28 which has a leading edge 34 which extends from a blade root 30 to a blade tip 40. The blade 20 is integrated with the hub 36, i.e. such the blades 20 are integrally formed as a monolithic component with the fan hub 36 to form an IBR fan. The nose cone 22 of the engine may be fastened to an upstream end of the fan hub 36 by a plurality of fasteners 29.

When the radius of the leading edge 30 on the hub 36 is reduced while the radius of the blade tip at 40 is maintained, the flow area (FA) of the fan 20 is increased thus reducing the specific flow (SF). As seen in FIG. 2, the gaspath through the fan 12 is defined by the annular area between the hubs 30 and the tips 40 of the fan blades 20. The radius of the fan hub (R_(HUB)), measured at the leading edge 34 of the blade 20, defines the radially inner gaspath boundary and the radius of the blade tip (R_(TIP)), also measured at the leading edge 34, defines the radially outer gaspath boundary. The specific flow of the fan 12 is therefore defined as the mass flow (MF) of air entering the leading edge of the fan 12, divided by the flow area (FA) at the fan leading edge, normal to the engine axis 13.

The hub to tip ratio of the IBR fan 12 is defined as the ratio of the radius of the fan hub (R_(HUB)) at the leading edge divided by the radius of fan blade tip (R_(TIP)) at the leading edge. As shown in FIG. 2, these radii are is measured from the engine centerline axis 13.

Thus, specific flow is determined as follows: SF=MF/FA, where SF is the specific flow, MF is the mass flow, and FA the flow area. Reduction of this SF of the fan is desirable as a reduced SF helps to improve the overall aerodynamic efficiency of the fan because of the lower air velocity.

A reduction in the hub to tip ratio (R_(HUB)/R_(TIP)) will therefore also cause a reduction in the specific flow (SF) of the fan. Alternatively, the radius of both the hub 36 and the blade tip 40 can be reduced while retaining the same specific flow SF. However, the ratio of the hub to tip radii is preferably reduced. Accordingly, the present IBR fan 12 has a ratio of the hub radius to the tip radius, i.e. R_(HUB)/R_(TIP), which is at least less than 0.29. In a particular embodiment, the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is between about 0.25 and about 0.29. In a further particular embodiment, the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is less than or equal to 0.25.

The advantage of a lower tip radius is a smaller diameter fan and therefore a lighter weight engine. Lowering the hub leading edge radius also changes the flow angle of the airstream, and the resulting rearward sweep in the lower portion of the fan blade airfoils 28 improves performance by reducing the leading edge velocities through the sweep effect and also draws flow towards the hub 36 which helps to reduce flow separation that the blade root.

The advantage of using the integrally bladed rotor (IBR) fan 12 is its reduced weight compared to a traditional detachable bladed rotor. The machining of an IBR fan 12 with such a low hub/tip ratio is made difficult by the lack of space between the blades 20, particularly at the blade roots 30 since the gap between the blades is much narrower the smaller the radius of the fan.

However, in one particular method of manufacturing the IBR fan 12 described herein, it has been found that by first machining a root stub 44 on the hub 36, the lower hub radius, and more particularly the low hub to tip radius ratios described above, can be obtained because it is easier to access the radial gap between adjacent blades 20 with machine tools. The blade airfoils 28 may then be fixed to the root stubs 44 be welded by Linear Friction Welding (LFW), for example, along the joint line 42 as shown on the blade 20 in FIG. 2. It has been contemplated that alternative methods may also be used, such as forming a root stub 44 only for every alternate blade, while machining the full blade 20 between each alternate root stub. This would allow sufficient access for machine tools between two alternate full blades, to machine around the around the remaining root stub.

Thus, a low-weight fan 12 as described herein is achieve, because of its integrated bladed rotor construction, and which provides a hub to tip radius ratio of at least less than 0.29, and more particularly between 0.25 and 0.29, and more particularly still a hub to tip radius ratio of 0.25 or less.

The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described while still falling within the scope of the appended claims, which define the present invention. Such modifications will be apparent to those skilled in the art, in light of a review of this disclosure. 

The invention claimed is:
 1. A fan for a turbofan gas turbine engine, the fan comprising a rotor hub and a plurality of fan blades arranged in a single blade row on the rotor hub, the fan blades radially extending from and being integral with the hub to form an integrally bladed rotor, each of the fan blades of the single blade row having a leading edge, a hub radius (R_(HUB)) which is the radius of the leading edge at the hub relative to a centerline of the fan, and a tip radius (R_(TIP)) which is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan, and wherein the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is than 0.29.
 2. The fan as defined in claim 1, wherein the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is less than about 0.25.
 3. The fan as defined in claim 1, wherein the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is between 0.25 and 0.29.
 4. The fan as defined in claim 1, wherein the fan preform includes root stubs disposed on the hub at positions corresponding to at least alternate ones of said fan blades, the root stubs being first formed on the hub prior to blades being fastened thereto.
 5. The fan as defined in claim 4 wherein the root stubs formed on the preform have airfoils welded thereto to provide the integrally bladed rotor.
 6. The fan as defined in claim 5 wherein the airfoils are linear-friction-welded to the respective root stubs.
 7. The fan as defined in claim 4 wherein all of the fan blades are first formed as root stubs on the hub.
 8. A method of manufacturing an integrally bladed rotor fan for a turbofan gas turbine engine, comprising: forming a rotor hub preform defining a hub radius and having at least a number of root stubs circumferentially spaced apart on a periphery of the rotor hub preform and aligned in a single blade row thereon; providing blade airfoils having a length such that a ratio of the hub radius to a tip radius of the blade airfoils, once mounted to the hub, is less than 0.29; and subsequently fastening the blade airfoils to the root stubs to form fan blades integrally formed with the hub—resulting in an integrally bladed rotor fan having the fan blades thereof arranged in the single blade row and having a hub to tip radius ratio of less than 0.29.
 9. The method as defined in claim 8, wherein the step of selecting further comprises selecting the length of the blade airfoils to define a ratio of the hub radius to the tip radius of the blade airfoils that is between 0.25 and 0.29.
 10. The method as defined in claim 9, wherein the step of selecting further comprises selecting the length of the blade airfoils to define a ratio of the hub radius to the tip radius of the blade airfoils that is less than about 0.25.
 11. The method as defined in claim 8, wherein circumferentially alternate ones of said fan blades are formed integrally with the hub perform without root stubs, leaving alternate root stubs on the hub preform to provide access for machine tools between the circumferentially alternate ones of said fan blades.
 12. The method as defined in claim 8, wherein the step of fastening further comprises welding the blade airfoils to the root stubs using Linear Friction Welding.
 13. A turbofan gas turbine engine comprising a fan upstream of at least one compressor, the fan having a rotor hub and a plurality of substantially radially extending fan blades integral with the rotor hub to form an integrated bladed rotor, the fan blades being arranged in a single blade row on the rotor hub, each said fan blade of the single blade row having an airfoil defining a leading edge and defining a tip radius (R_(TIP)) which is the radius of a tip of the fan blade at the leading edge, the rotor hub defining a hub radius (R_(HUB)) which is the radius of the hub at the blade leading edge, and wherein a ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is less than 0.29.
 14. The turbofan gas turbine engine as defined in claim 13, wherein the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is less than about 0.25.
 15. The turbofan gas turbine engine as defined in claim 13, wherein the ratio of the hub radius to the tip radius (R_(HUB)/R_(TIP)) is between 0.25 and 0.29. 